This invention relates generally to gas turbine engine nozzles and more particularly, to methods and apparatus for assembling gas turbine engine nozzles.
Gas turbine engines include combustors which ignite fuel-air mixtures which are then channeled through a turbine nozzle assembly towards a turbine. At least some known turbine nozzle assemblies include a plurality of nozzles arranged circumferentially and configured as doublets. At least some known turbine nozzles include more than two circumferentially-spaced hollow airfoil vanes coupled by integrally-formed inner and outer band platforms. More specifically, the inner band forms a radially inner flowpath boundary and the outer band forms a radially outer flowpath boundary. Other known turbine nozzles are mounted in a cantilever arrangement wherein the inner band is moveable radailly and axially, and the outer band is constrained at forward and aft hooks.
Forming the turbine nozzle with greater than two integrally-formed airfoil vanes facilitates improving durability and reducing leakage in comparison to turbine nozzles which include only one or two airfoil vanes. Accordingly, at least some known turbine nozzles include at least one airfoil vane positioned between a pair of circumferentially outer airfoil vanes. However, during operation, temperature gradients and aerodynamic loading may result in thermal stresses and thermal chording at an interface between the airfoil vanes and the outer band. More specifically, higher stresses may be induced into the outer airfoil vanes than the vanes positioned between the outer airfoil vanes. Over time, the local stresses induced to the turbine nozzle may cause premature failure of the turbine nozzle.
To facilitate reducing the effects of thermal gradients and aerodynamic loading, within at least some known turbine nozzles, a compound radii fillet is formed between each airfoil vane and the outer band. However, because at least some known turbine nozzles are designed with low aerodynamic convergence to permit an easy passage for airfoil cooling and to pass cooling and purge air for the high pressure turbine/low pressure turbine rotor cavities. Thus, extending compound radii fillets along the airfoil vanes may undesirably reduce aerodynamic convergence through the turbine nozzle. Furthermore, in extreme circumstances, the reduced aerodynamic convergence may cause the nozzle aerodynamic throat to shift forward from the nozzle trailing edge, thus resulting in an unstable aerodynamic environment.